This disclosure relates to a cooling passage for an airfoil.
Turbine blades are utilized in gas turbine engines. As known, a turbine blade typically includes a platform having a root on one side and an airfoil extending from the platform opposite the root. The root is secured to a turbine rotor. Cooling circuits are formed within the airfoil to circulate cooling fluid, such as air. Typically, multiple relatively large cooling channels extend radially from the root toward a tip of the airfoil. Air flows through the channels and cools the airfoil, which is relatively hot during operation of the gas turbine engine.
Some advanced cooling designs use one or more radial cooling passages that extend from the root toward the tip. Typically, the cooling passages are arranged between the cooling channels and an exterior surface of the airfoil. The cooling passages provide extremely high convective cooling.
The Assignee of the present disclosure has discovered that in some cooling designs the airfoil is overcooled at the base of the airfoil near the platform. It is believed that strong secondary flows, particularly on the suction side, force the migration of relatively cool fluid off the end wall and onto the suction side of the blade. This results in relatively low external gas temperatures. Internally, the coolant temperature is relatively cool as it has just entered the blade. The high heat transfer coefficients provided by the cooling passage in this region are undesirable as it causes overcooling of the external surface and premature heating of the coolant air.
What is needed is a cooling passage that provides desired cooling of the airfoil.